The present invention relates to the specific field of turbomachines and it relates more particularly to the problem posed by assembling a combustion chamber made of a composite material of the ceramic matrix composite (CMC) type in the metal chamber casings of a turbomachine.
Conventionally, in a turbojet or a turboprop, the high pressure turbine, in particular its inlet nozzle (HPT nozzle), the combustion chamber, and the inner and outer shells (or casings) of said chamber are all made out of the same material, generally a metal. Nevertheless, under certain particular conditions of use implementing particularly high combustion temperatures, a metal chamber turns out to be completely unsuitable from a thermal point of view and it is necessary to make use of a chamber that is based on high temperature composite materials of the CMC type. However, difficulties of implementation and materials costs mean that such materials are generally restricted to being used for the composite chamber itself, with the high pressure turbine inlet nozzle and the inner and outer shells of the chamber then still being made more conventionally out of metal materials. Unfortunately, metals and composites have coefficients of thermal expansion that are very different. This gives rise to particularly awkward problems of connection with the inner and outer shells of the combustion chamber and of interface at the nozzle at the inlet to the high pressure turbine.
The present invention mitigates those drawbacks by proposing a mounting for the combustion chamber in the casings with the ability to absorb the displacements induced by the various coefficients of expansion of those parts. An object of the invention is thus to propose a mounting which makes the best use of the existing characteristics of the combustion chamber.
These objects are achieved by a turbomachine comprising inner and outer annular shells of metal material containing in a gas flow direction F: a fuel injector assembly, an annular combustion chamber of composite material and having a longitudinal axis, and an annular nozzle of metal material and forming the fixed-blade inlet stage of a high pressure turbine, wherein said composite material combustion chamber is held in position between said inner and outer metal annular shells by a plurality of flexible tongues, first ends of said tongues being interconnected by a metal ring fixed securely to each of said inner and outer metal annular shells by first fixing means, and second ends being fixed by second fixing means to a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells.
With this particular structure for the fixed connection, the various kinds of wear due to contact corrosion in prior art systems can be avoided. The use of a ring made of composite material to provide sealing of the stream also makes it possible to keep the initial structure of the chamber intact. In addition, the presence of flexible metal tongues replacing the traditional flanges gives rise to a saving in mass that is particularly appreciable. In addition to being flexible, these tongues make it easy to accommodate the expansion difference that appears at high temperatures between metal parts and composite parts (by accommodating the displacements due to expansion) while still ensuring that the combustion chamber is properly held and well centered in the annular shell.
The first and second fixing means are preferably constituted by a plurality of bolts.
In an advantageous embodiment in which each of said metal annular shells is made up of two portions, said metal ring interconnecting said first ends of said metal fixing tongues is mounted between connecting flanges of said two portions. In an alternative embodiment, said metal ring can be fixed directly to said annular shell by fixing means.
Depending on the intended embodiment, said first ends of the fixing tongues can either be fixed by brazing to said metal ring, or else they can be formed integrally with said metal ring.
In a preferred embodiment, said composite ring is brazed onto a downstream end of the combustion chamber. In an alternative embodiment, the composite ring is sewn onto the downstream end. In another embodiment, the composite ring is implanted on the downstream end.
Said composite ring includes a determined portion forming a bearing plane for a sealing gasket (advantageously of the circular xe2x80x9cspring bladexe2x80x9d gasket type) ensuring that the stream of gas between said combustion chamber and said nozzle is sealed. Said determined portion is preferably an end portion of said composite ring.